A sealing apparatus for an axial flow turbomachine is disclosed. An example axial flow turbomachine includes band segments bordering a flow path, a first spline seal disposed at a first angle within a first slot of the band segments, a second spline seal disposed at second angle within a second slot of the band segments, a Z-seal disposed across a gap between the first slot and the second, the Z-seal positioned at least partially within the first slot and the second slot. The Z-seal includes a first segment disposed at the first angle at least partially within the first slot, the first segment to contact the first spline seal, a second segment disposed at the second angle at least partially within the second slot, a portion of the second segment in connection with a portion of the first segment, the second segment to contact the second spline seal.

A center plug includes an inner skin, the inner skin extending along an axial centerline; an outer skin positioned radially outside the inner skin; a forward bulkhead connected to and extending radially outward from the inner skin; an aft bulkhead connected to and extending radially outward from the inner skin; and an acoustic panel, the acoustic panel including a first sheet having a first plurality of perforated walls and a first plurality of non-perforated walls, a second sheet having a second plurality of perforated walls and a second plurality of non-perforated walls, the first sheet being sandwiched together with the second sheet to form an N-shaped structure having a combined plurality of perforated walls and a combined plurality of non-perforated walls.

The present technique presents a turbomachine component having an airfoil e.g. a vane of a gas turbine. The airfoil wall defines an internal space which includes a first and a second cooling channels having a first and a second impingement inserts, that define a first main and a first peripheral flow channels in the first cooling channel and a second main and a second peripheral flow channels in the second cooling channel, respectively. Impingement jets ejected from the main flow channels via impingement holes of the corresponding impingement inserts are received in the corresponding peripheral flow channels. A channel connecting conduit conducts a flow of the cooling air from the first cooling channel to the second cooling channel. The channel connecting conduit includes an inlet connected to an outlet of the first cooling channel, and an outlet connected to an inlet of the second cooling channel.

Disclosed is an outer air seal, having: an axial member, the axial member extending axially from an axial front end to an axial aft end, and extending radially from a radial inner surface to a radial outer surface; a radial flange extending radially from the radial outer surface of the axial member to a radial outer tip, and extending axially from an axial front surface to an axial aft surface; and a first kerf slot defined through the axial member from the axial front end to the axial aft end and from the radial inner surface to the radial outer surface, and through the radial flange from the axial front surface to the axial aft surface, wherein a radial top end of the first kerf slot is radially spaced apart from the radial outer tip of the radial flange.

The invention relates to a turbine assembly comprising: a plurality of ring sectors forming a ring; a support structure comprising a shroud from which there project radial flanges which serve to hold latching tabs of each ring sector; a stator positioned downstream of said ring and comprising a blade provided with a radially outer platform and extending axially opposite said ring; an annular space being defined between the ring, the support structure and the platform, this annular space having passing through it a leakage air current, the assembly comprising a hollow ring which occupies said annular space and which is shaped so as to collect, channel and expel said leakage air current, more specifically to channel and expel it into the radially inner region of said annular space.

A bladed disc system for a turbine engine having a disk portion and a plurality of blade portions which are associated with a stator section and an intercavity sealing portion, disc portion shaped such that blade portions are able to fit within firtree slot in disc portion, blade portion having aerofoil section and root section, aerofoil section having portion shaped such that they extend proximate to intercavity sealing portion, disc portion extending from portion that connects with drum to outer edge at which blade portions are connected with disc portion having width transition region in which thickness of disc increases from point at which disc connects to drum to outer edge at which it holds blade portions, and wherein width transition region has curved width transition region with radius r, and an overhanging portion which extends into the intercavity spacing between the width transition region and the intercavity sealing portion.

A bladed disc system for a turbine engine, the bladed disk portion comprising a disk portion and a plurality of blade portions, the disc portion being shaped such that the blade portions are able to fit within fir tree slot in the disc portion, the blade portion comprising an aerofoil section and a root section, with the root section comprising a fir tree profile and a skirt portion and wherein the skirt portions of adjacent blades form an opening that has a maximum separation of between 1-50% of the maximum skirt opening width.

A structural and/or acoustic panel including an inner skin, an outer skin, a honeycomb structure and at least one sealing flange. The sealing flange has a U-shaped section directed toward the honeycomb structure and is clamped with the honeycomb structure between the inner and the outer skins. A method for manufacturing such a panel by soldering the component elements thereof is also disclosed.

A turbine engine that includes an engine core, an inner cowl, an outer cowl and an accessory gearbox. The engine core includes at least a compressor section, a combustion section, and a turbine section in axial flow arrangement. The accessory gearbox is operably coupled to the engine core and includes a first portion and a second portion.

A discharge grille of an intermediate casing of a turbomachine includes a peripheral edge and a seal mounted on the peripheral edge via an attachment system, the seal ensuring the sealing between a first turbomachine member and a second member of the turbomachine, and surrounding the discharge grille of the intermediate casing, the seal having a U-shaped cross section and includes a base plate configured to bear against the first member, a sealing lip configured to bear against the second member through deformation of the sealing lip to ensure air tightness and fire proofing through contact.